1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine airfoil with cooling circuits.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as an industrial gas turbine (IGT) engine, stator vanes and rotor blades (both airfoils) used in the turbine air cooled by passing pressurized air from the compressor through a complex shaped internal cooling circuit within the airfoil. The efficiency of the engine can be increased by allowing for a higher gas flow temperature entering the turbine, or by using less cooling air to cool the airfoils.
FIG. 1 shows a prior art first stage rotor blade external pressure profile. The forward region of the pressure side surface experiences a high hot gas static pressure while the entire suction side of the airfoil is at a much lower hot gas static pressure than on the pressure side. On the pressure side of the airfoil, the pressure profile increases from the leading edge to a high point just downstream from the leading edge, and then drops in the direction toward the trailing edge. On the suction side of the airfoil, the pressure profile shows the highest pressure to be at the leading edge, and a significant drop in pressure just past the leading edge to a low point at about mid-chord length with the pressure increasing toward the trailing edge. Most of the pressure on the suction side is below the lowest pressure found on the pressure side as indicated in FIG. 1.
U.S. Pat. No. 6,988,872 B2 issued to Soechting et al on Jan. 24, 2006 and entitled TURBINE MOVING BLADE AND GAS TURBINE shows this serpentine cooling circuit is reproduced in FIG. 2 of the present application and shows a prior art rotor blade with a (1+3+3) serpentine flow cooling circuit for the first stage blade. For the first forward flowing triple pass (also referred to as a three-pass) serpentine flow cooling circuit used in the airfoil mid-chord region, the cooling air flows in the forward direction and discharges into the high hot gas side pressure section of the pressure side. In order to satisfy the backflow margin criteria, a high cooling supply pressure is needed for the prior art cooling circuit of FIG. 2. The high pressure requirement induces a high leakage flow. Since the last up-pass of the triple pass serpentine cavities provide film cooling air for both sides of the airfoil, in order to satisfy the back flow margin criteria for the pressure side film row, the internal cavity pressure must be approximately 10% higher than the pressure side hot gas side which will result in an over-pressuring of the airfoil suction side film cooling holes. For the second aft flowing triple pass serpentine flow cooling design used in the airfoil mid-chord region, the cooling air flows forward and discharges into the airfoil trailing edge to provide cooling for the airfoil trailing edge region.
Another prior art reference, U.S. Pat. No. 4,820,122 issued to Hall et al on Apr. 11, 1989 and entitled DIRT REMOVAL MEANS FOR AIR COOLED BLADES discloses an internally cooled turbine blade in FIG. 1 of this patent with a two passages (28 and 30 in this patent) that are adjacent and parallel and form the first and second legs of a serpentine flow passage, but share a third leg in which the two passages merge into the third leg. These two passages are not separate from each other. Also, the flow direction in this patent is from trailing edge to leading edge, the opposite direction of the hot gas flow through the turbine.
U.S. Pat. No. 6,220,817 B1 issued to Durgin et al on Apr. 24, 2001 and entitled AFT FLOWING MULTI-TIER AIRFOIL COOLING CIRCUIT discloses a turbine airfoil with two serpentine flow cooling passages arranged one above the other in which an outer inlet channel (40b in this patent) extends to the blade tip and then into a first serpentine flow circuit in the upper portion of the blade and a inner inlet channel (40a in this patent) than extends about halfway up and then flows into a second serpentine flow circuit in the lower portion of the blade. The two serpentine flow circuits are separated by ribs so that no cross-flow of cooling air occurs, and both discharge through blade tip cooling holes. U.S. Pat. No. 5,591,007 issued to Lee et al on Jan. 7, 1997 and entitled MULTI-TIER TURBINE AIRFOIL discloses a similar cooling circuit arrangement to the above Durgin patent.
U.S. Pat. No. 5,403,159 issued to Green et al on Apr. 4, 1995 and entitled COOLABLE AIRFOIL STRUCTURE discloses a turbine airfoil with a cooling circuit with a passage (82 to 86 in this patent) that includes a rib with a curved top end that breaks the channel into two separate channels. However, the channel is not two separate serpentine channels as is the applicant's present invention.
It is therefore an object of the present invention to provide for a new serpentine flow cooling circuit that can be used in a rotor blade, especially for the mid-chord and trailing edge regions of the blade, in which the above described disadvantages are reduced or eliminated.